Gas turbine engine aft bearing arrangement

ABSTRACT

An example gas turbine engine includes a turbine and first and second spools coaxial with one another. The first spool is arranged within the second spool and extends between forward and aft ends. The aft end extends axially beyond the second spool and supports the turbine. A housing is arranged downstream from the turbine. First and second bearings are mounted to the aft end of the first spool and supported by the housing portion.

This application is a continuation application of U.S. application Ser.No. 14/033,652, filed on Sep. 23, 2013, which is a continuation in partapplication of U.S. application Ser. No. 13/567,178, filed on Aug. 6,2012, which is a divisional application of U.S. application Ser. No.13/364,502, filed on Feb. 2, 2012, which claims priority to U.S.Provisional Application No. 61/593,050, filed on Jan. 31, 2012.

BACKGROUND

This disclosure relates to a bearing arrangement for a gas turbineengine.

A typical jet engine has two or three spools, or shafts, that transmittorque between the turbine and compressor sections of the engine. Eachof these spools is typically supported by two bearings. One bearing, forexample, a ball bearing, may be arranged at a forward end of the spooland be configured to react to both axial and radial loads. Anotherbearing, for example, a roller bearing or journal bearing may bearranged at the aft end of the spool and be configured to react only toradial loads. This bearing arrangement typically fully constrains theshaft except for rotation, and axial movement of one free end ispermitted to accommodate engine axial growth.

SUMMARY

A bearing hub for a gas turbine engine according to an exemplaryembodiment of this disclosure, among other possible things includes aradial to axial translation flange arm extending outward from an apex ofthe unitary structure, a translation flange extending outward from saidtranslation flange arm, a spring arm connected to the apex forconnecting the bearing hub to a canted annular flange.

In a further embodiment of the foregoing bearing hub, the translationflange arm extends axially aftward from the apex of the unitarystructure.

In a further embodiment of the foregoing bearing hub, the spring armcomprises at least a first flex point, a second flex point, and a thirdflex point, and a stiffness of each of the flex points is configured todetermine an amount of radial vibrations translated to axial vibrationsby the bearing hub.

In a further embodiment of the foregoing bearing hub, the first andsecond hub walls are inclined radially inward from an annular apex, anda first and second bearing are respectively supported by the first andsecond walls opposite the apex.

In a further embodiment of the foregoing bearing hub, a focal node ofradial vibrations of the bearing hub is the first bearing.

In a further embodiment of the foregoing bearing hub, the spring arm isrigidly connected to the apex.

A gas turbine engine according to an exemplary embodiment of thisdisclosure, among other possible things includes a fan, a compressorsection fluidly connected to the fan, the compressor including a firstcompressor section and a second compressor section, a combustor fluidlyconnected to the compressor section, a turbine section fluidly connectedto the combustor, the turbine section including a first turbine sectioncoupled to the first compressor section via a shaft, a second turbinesection, first and second hub walls integrally formed with one anotherto provide a unitary structure, a radial to axial translation flange armextending outward from an annular apex of the unitary structure, atranslation flange extending outward from the translation flange arm,turbine exhaust case arranged downstream from the second turbine sectionand supporting the annular apex, and a spring arm connecting the apex toa canted annular flange of the turbine exhaust case.

In a further embodiment of the foregoing gas turbine engine, thetranslation flange arm extends axially aftward from the apex of theunitary structure.

In a further embodiment of the foregoing gas turbine engine, thetranslation flange is received in an annular cavity supported by thecanted annular flange.

In a further embodiment of the foregoing gas turbine engine, the annularcavity includes an axial vibration damper.

In a further embodiment of the foregoing gas turbine engine, the axialvibration damper includes at least a first wire mesh structure disposedbetween the translation flange and a first wall of the annular cavity.

In a further embodiment of the foregoing gas turbine engine, the axialvibration damper includes at least a second wire mesh structure disposedbetween the translation flange and a second wall of the annular cavity.

In a further embodiment of the foregoing gas turbine engine, the axialvibration damper includes at least a first seal defining a dampingannulus within the annular cavity.

A further embodiment of the foregoing gas turbine engine includes adamping fluid disposed within the damping annulus.

In a further embodiment of the foregoing gas turbine engine, the dampingfluid is damping oil.

In a further embodiment of the foregoing gas turbine engine, the axialvibration damper includes at least a second seal further defining thedamping annulus.

In a further embodiment of the foregoing gas turbine engine, the seal isone of an elastomeric O-ring seal and a piston ring.

In a further embodiment of the foregoing gas turbine engine, the secondturbine section is a low pressure turbine, and the low pressure turbineis configured to have a pressure ratio that is greater than about 5:1.

A method for damping vibrations in a bearing hub according to anexemplary embodiment of this disclosure, among other possible thingsincludes converting radial vibrations in the bearing hub to axialvibrations using a spring arm and a radial to axial vibrationtranslation flange, damping axial vibrations of the radial to axialvibration translation flange using an axial vibration damper.

A further embodiment of the foregoing method, includes the step ofadjusting a level of vibrational damping by adjusting a stiffness of thespring arm.

The foregoing features and elements may be combined in any combinationwithout exclusivity, unless expressly indicated otherwise.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings.

FIG. 1 schematically illustrates an example gas turbine engine.

FIG. 2 is another schematic view of the gas turbine engine illustratingan example bearing arrangement.

FIG. 3 is a more detailed schematic view of an aft bearing arrangementillustrated in FIG. 2.

FIG. 4 schematically illustrates a more detailed alternative example ofthe aft bearing arrangement illustrated in FIG. 2.

FIG. 5 schematically illustrates a first damping annulus for thealternative example aft bearing arrangement of FIG. 4.

FIG. 6 schematically illustrates a second damping annulus for thealternative example aft bearing arrangement of FIG. 4.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure or housing 58 via severalbearing systems 38. It should be understood that various bearing systems38 at various locations may alternatively or additionally be provided.The housing 58 includes first, second, third, and fourth housingportions 58A, 58B, 58C, 58D. The third and fourth housing portions 58C,58D respectively correspond to a mid-turbine frame and a turbine exhaustcase.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. The mid-turbine frame 58C of theengine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 58C supports one or more bearing systems 38 in the turbine section28. The turbine exhaust case 58D is arranged downstream from the lowpressure turbine 46 and may support one or more bearing systems as well.The inner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A, whichis collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 is, in one example, a high-bypass geared aircraft engine.In a further example, the engine 20 bypass ratio is greater than aboutsix (6), with an example embodiment being greater than ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about 5. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Temperature ambientdeg Rankine)/518.7)^0.5]. The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second.

An example bearing arrangement for the gas turbine engine 20 isschematically depicted in FIG. 2. The low speed spool 30 extends axiallybetween forward and aft ends, with the aft end extending beyond the highspeed spool 32. The high speed spool 32 is supported for rotation byfirst and second bearings 38A, 38B. In one example, the first bearing38A is of a type that may react to both axial and radial loads, such asa ball bearing. The second bearing 38B is of a type that reacts toradial loads, such as a roller bearing or a journal bearing. The lowspeed spool 30 is supported by first, second and third bearings 38C,38D, 38E. The first bearing 38C is of a type that reacts to both axialand radial loads. The second and third bearings 38D, 38E are of a typethat reacts to radial loads. The second and third bearings 38D, 38Esupport an aft end of the low speed spool 30.

The low speed spool 30 has a higher length/diameter (L/D) ratio than thehigh speed spool 32. From a rotor dynamics standpoint, a shaft willreach a critical speed of instability at a lower speed as the L/D ratiogets larger. Providing at least two bearings at the aft end of the lowspeed spool 30 increases the critical speed of the low speed spool 30,which enables higher overall engine speeds and lower weight therebyallowing the engine 20 to be faster and smaller for a given level ofthrust.

FIG. 3 illustrates a more detailed view of the engine 20 shown in FIG.2. The second bearing 38B supporting the low speed spool 30 is arrangedwithin a first bearing compartment 60. The first bearing compartment 60is provided by first and second walls 62, 64, which are sealed relativeto the third housing portion 16C and the second spool 30.

The low pressure turbine 46 includes a low pressure turbine rotor hub 66secured to the low speed spool 30. The low pressure turbine rotor hub 66supports multiple low pressure turbine blades 68 in one example. Lowpressure turbine stator vanes 70 are provided between the low pressureturbine blades 68 and supported by the housing 16. The low pressureturbine rotor hub 66 is canted in an aft direction, which accommodates asecond bearing compartment 72. The second and third bearings 38D, 38Eare arranged within the second bearing compartment 72, which is providedby first and second walls 80, 82 and a cover 84, for example. The cover74 is removably secured over the aft end and encloses the second andthird bearings 38D, 38E.

In the example, the turbine exhaust case 58D includes a radiallyrearward canted annular flange 76, which removably supports a hub 74secured to the flange 76 by fasteners 78. The hub 74 includes first andsecond hub walls 86, 88 canted toward one another in opposite directionsand adjoining one another at an annular apex 89 provided near the flange76 in the example shown. In the example shown, the first and second hubwalls 86, 88 provide an integrated, unitary structure. Each of the firstand second hub walls 86, 88 supports one of the second and thirdbearings 38D, 38E. The aft-canted low pressure turbine rotor hub 66accommodates at least the second bearing 38D and a portion of the firsthub wall 86 is arranged radially beneath the low pressure turbine 46such that axial length need not be added to the low speed spool 30. Thesecond bearing 38D is arranged axially forward of an aft side 98 of alast rotor stage 96.

In one example, the second and third bearings 38D, 38E are spaced apartfrom one another a span 90 that is approximately 4-12 inches (10-30 cm),for example. In one example, the first hub wall is oriented at a firstangle 92 of between about 30° and about 60°, and the second hub wall 88is oriented at a second angle 94 of between about 30° and about 60°. Theannular flange 76 is oriented relative to the first hub wall 86 at athird angle 95 of between about 0° and about 30°, for example. In oneexample, the first angle 90 is about 45° and the third angle 95 is about0°. The above values are exemplary for one example engine design.

The multiple turbine rotors include first, second and third turbinerotors 69A, 69B, 69C. The third turbine rotor is part of the last rotorstage 96. The first turbine rotor 69A corresponds to a forward-mostrotor stage. The second turbine rotor 69B is arranged axially betweenthe first and third rotors 69A, 69C. The low pressure turbine rotor hub66 is mounted on the low speed spool 30 and is secured to the secondturbine hub 69B for supporting the low pressure turbine 46. In oneexample, the bearing hub 74 includes a moment stiffness of about80,000,000 in-lb/rad (9,144,000 cm-kg/rad), for example, and a lateralstiffness of about 5,000,000 lb/in (886,000 kg/cm).

With continued reference to FIGS. 1-3, and with like numerals indicatinglike elements, FIG. 4 schematically illustrates a more detailedalternative example of the aft bearing arrangement illustrated in FIG.2. The alternate aft bearing arrangement 100 of FIG. 4 includes anannular cavity 112 defined by annulus walls 110. The annulus walls 110are supported by the radially rearward canted annular flange 76.

A radial to axial translation flange arm 120 extends from the apexdefined by the first and second hub walls 86, 88 and supports a radialto axial translation flange 122. Also connecting the apex of the firstand second hub walls 86, 88 to the radially rearward canted annularflange 76 is a spring arm 130. The spring arm 130 has multiple flexpoints 132 that cooperate to convert radial vibrations 140 in the aftbearing arrangement 100 into axial vibrations 150 in the radial to axialtranslation flange 122.

During operation of the turbine engine, the aft bearing arrangement 100vibrates about the second bearing 38D connecting the aft bearingarrangement 100 to the low speed spool 30. Due to the design of the aftbearing arrangement 100, the second bearing 38D is the focal node of theaft bearing radial vibrations 140 (illustrated by dashed lines). Theillustrated radial vibrations 140 are exaggerated for illustrativeeffect.

The spring arm 130 operates cooperatively with the radially rearwardcanted annular flange 76 to convert the radial vibrations 140 into axialvibrations 150 (illustrated by dashed lines) by flexing at the flexpoints 132 and along the longitudinal cylinders comprising the springarm 130. The axial vibrations 150 have a focal node 152 centered in thesecond hub wall 88. In alternate examples, the focal node of the axialvibrations 150 can be located anywhere radially aligned with theillustrated focal node 152, as dictated by the design of the spring arm130.

The magnitude of the axial vibrations 150 and the amount of translationfrom the radial vibrations 140 to the axial vibrations 150 is directlyrelated to the stiffness of the spring arm 130. One of skill in the art,having the benefit of this disclosure, can adjust the amount of radialvibrations 140 translated into axial vibrations 150 by adjusting thestiffness of the spring arm 130. In a practical embodiment, the springarm 130 stiffness is adjusted during engine design and a spring arm 130having a fixed spring stiffness of the determined stiffness is utilizedin the actual assembly.

Once the radial vibrations 140 are translated into axial vibrations 150,the vibrations 150 are damped via an axial vibration damper positionedwithin the annular cavity 112. With continued reference to FIGS. 1-4,and with like numerals indicating like elements, FIG. 5 illustrates afirst example axial vibration damping arrangement 200 positioned withinthe annular cavity 112. The radial to axial translation flange 122 isreceived in the annular cavity 112 and a pair of seals 160, 162 createsa damping annulus 164 within the annular cavity 112. In one example, theseals 160, 162 are elastomeric O-ring seals. In an alternate example,the seals 160, 162 are a non-elastomeric seal, such as a piston ring. Afluid, such as oil, is received within the damping annulus 164 anddampens the axial vibrations 150 by axially compressing anddecompressing as the radial to axial translation flange 122 vibrates.

With continued reference to FIGS. 1-4, and with like numerals indicatinglike elements, FIG. 6 illustrates an alternate axial damping arrangement300 for the aft bearing arrangement illustrated in FIG. 2. As with theaxial damping arrangement 200 of FIG. 5, the radial to axial translationflange 122 is received in the annular cavity 112 defined by the annuluswalls 110. Instead of a fluid filled damping annulus 164 defined by apair of seals, the radial to axial flange 122 is received between a pairof metal mesh dampers 170, 172. As the radial to axial flange 122vibrates axially, the metal mesh dampers 170, 172 compress anddecompress, thereby damping the vibration of the aft bearingarrangement, and reducing the negative impact of vibrations. The amountof damping provided by the metal mesh dampers 170, 172 can also beadjusted by adjusting the stiffness of the metal mesh dampers 170, 172.

With continued reference to FIGS. 1-6, utilization of the aft bearingarrangement 100 allows the trunnion bearings to be spaced closer thanconventional trunnion bearings, as the axial vibration damping has agreater damping effect. Similarly, while the above examples aredescribed with regards to an aft bearing arrangement, the disclosedstructure can be adapted to other radial damping bearing arrangements byone skilled in the art having the benefit of this disclosure in order toreplace the radial damping with axial damping, or, alternatively, toaugment existing radial damping.

Although example embodiments have been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A bearing hub for a gas turbine enginecomprising: first and second hub walls integrally formed with oneanother to provide a unitary structure; a radial to axial translationflange arm extending outward from an apex of the unitary structure; atranslation flange extending outward from said radial to axialtranslation flange arm; a spring arm connected to the apex forconnecting the bearing hub to a canted annular flange, the spring armincluding a plurality of angled flex points.
 2. The bearing hub of claim1, wherein the radial to axial translation flange arm extends axiallyaftward from said apex of said unitary structure.
 3. The bearing hub ofclaim 1, wherein the plurality of angled flex points comprises at leasta first flex point, a second flex point, and a third flex point, andwherein a stiffness of each of said first flex point, said second flexpoint and said third flex point is configured to control an amount ofradial vibrations translated to axial vibrations by said bearing hub. 4.The bearing hub of claim 1, wherein the first and second hub walls areinclined radially inward from an annular apex, and a first and secondbearing are respectively supported by the first and second wallsopposite the apex.
 5. The bearing hub of claim 4, wherein a focal nodeof radial vibrations of the bearing hub is the first bearing.
 6. Thebearing hub of claim 1, wherein said spring arm is rigidly connected tosaid apex.
 7. A gas turbine engine comprising: a fan; a compressorsection fluidly connected to the fan, the compressor section comprisinga first compressor section and a second compressor section; a combustorfluidly connected to the compressor section; a turbine section fluidlyconnected to the combustor, the turbine section comprising: a firstturbine section coupled to the first compressor section via a shaft; asecond turbine section; first and second hub walls integrally formedwith one another to provide a unitary structure; a radial to axialtranslation flange arm extending outward from an annular apex of theunitary structure; a translation flange extending outward from saidradial to axial translation flange arm; turbine exhaust case arrangeddownstream from the second turbine section and supporting the annularapex; and a spring arm connecting the annular apex to a canted annularflange of the turbine exhaust case, the spring arm including a pluralityof angled flex points.
 8. The gas turbine engine of claim 7, wherein theradial to axial translation flange arm extends axially aftward from saidannular apex of said unitary structure.
 9. The gas turbine engine ofclaim 7, wherein said translation flange is received in an annularcavity supported by the canted annular flange.
 10. The gas turbineengine of claim 9, wherein said annular cavity includes an axialvibration damper.
 11. The gas turbine engine of claim 10, wherein saidaxial vibration damper comprises at least a first wire mesh structuredisposed between said translation flange and a first wall of saidannular cavity.
 12. The gas turbine engine of claim 11, wherein saidaxial vibration damper comprises at least a second wire mesh structuredisposed between said translation flange and a second wall of saidannular cavity.
 13. The gas turbine engine of claim 10, wherein saidaxial vibration damper comprises at least a first seal defining adamping annulus within said annular cavity.
 14. The gas turbine engineof claim 13, further comprising a damping fluid disposed within saiddamping annulus.
 15. The turbine engine of claim 14, wherein saiddamping fluid is damping oil.
 16. The turbine engine of claim 13,wherein said axial vibration damper comprises at least a second sealfurther defining the damping annulus.
 17. The turbine engine of claim13, wherein said first seal is one of an elastomeric O-ring seal and apiston ring.
 18. A method for damping vibrations in a bearing hubcomprising the steps of: converting radial vibrations in said bearinghub to axial vibrations using a spring arm and a radial to axialvibration translation flange, wherein the spring arm includes aplurality of angled flex points; damping axial vibrations of said radialto axial vibration translation flange using an axial vibration damper.19. The method of claim 18, further comprising the step of adjusting alevel of vibrational damping by adjusting a stiffness of the spring arm.